U.S. patent application number 14/077475 was filed with the patent office on 2015-05-14 for cooling air temperature reduction using nozzles.
The applicant listed for this patent is Chad W. Heinrich, Stephen Erick Holland. Invention is credited to Chad W. Heinrich, Stephen Erick Holland.
Application Number | 20150128598 14/077475 |
Document ID | / |
Family ID | 53042472 |
Filed Date | 2015-05-14 |
United States Patent
Application |
20150128598 |
Kind Code |
A1 |
Heinrich; Chad W. ; et
al. |
May 14, 2015 |
COOLING AIR TEMPERATURE REDUCTION USING NOZZLES
Abstract
A converging-diverging nozzle that has particular application
for providing a cooling air flow to ring segments in a gas turbine
engine. The engine includes a turbine section that receives a hot
working gas. The turbine section includes at least one row of
vanes, at least one row of blades and a plurality of ring segments
forming at least one ring. The ring segments and the vanes are
mounted to a vane carrier, where the vane carrier includes a
cooling flow channel for each of the ring segments that receives an
air flow to cool the ring segments. A plug is provided in each
channel and has an internal bore shaped to define the
converging-diverging nozzle through which the air flows so as to
create a supersonic flow that reduces the temperature of the air
and thus provides more cooling for the same amount of air flow.
Inventors: |
Heinrich; Chad W.; (Oviedo,
FL) ; Holland; Stephen Erick; (Oviedo, FL) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Heinrich; Chad W.
Holland; Stephen Erick |
Oviedo
Oviedo |
FL
FL |
US
US |
|
|
Family ID: |
53042472 |
Appl. No.: |
14/077475 |
Filed: |
November 12, 2013 |
Current U.S.
Class: |
60/726 ;
415/116 |
Current CPC
Class: |
F05D 2240/12 20130101;
Y02T 50/60 20130101; F02C 7/18 20130101; Y02T 50/675 20130101 |
Class at
Publication: |
60/726 ;
415/116 |
International
Class: |
F02C 7/18 20060101
F02C007/18 |
Claims
1. A gas turbine engine comprising: a compressor section being
operable to produce a compressed gas; a combustion section in fluid
communication with the compressor section that receives the
compressed gas, said combustion section mixing the compressed gas
with a fuel and combusting the compressed gas and fuel mixture to
produce a hot working fluid; and a turbine section in fluid
communication with the combustion section, said turbine section
receiving the hot working fluid to produce mechanical power, said
turbine section including at least one row of vanes, at least one
row of blades and a plurality of ring segments forming at least one
ring, wherein the ring segments provide a sealing structure at an
end of the blades, said ring segments and said at least one row of
vanes being mounted to a vane carrier, said vane carrier including
a ring cooling channel for each ring segment that receives a
portion of the compressed gas to cool the ring segments, said
turbine section further including a plug positioned in each ring
cooling channel and having an internal flow channel, said internal
flow channel defining a converging-diverging nozzle through which
the compressed gas flows so as to create a supersonic flow that
reduces the temperature of the compressed gas.
2. The gas turbine engine according to claim 1 wherein the internal
flow channel includes a wide end section coupled to a converging
section.
3. The gas turbine engine according to claim 2 wherein the internal
flow channel includes a narrow throat section coupled to the
converging section at one end and a diverging section at an
opposite end.
4. The gas turbine engine according to claim 3 wherein the internal
flow channel includes a wide output end section at an output end of
the nozzle coupled to the diverging section where the supersonic
flow occurs.
5. The gas turbine engine according to claim 1 wherein the turbine
section includes four rows of vanes, four rows of blades and four
rows of rings.
6. The gas turbine engine according to claim 5 wherein the ring
cooling channels are only provided for the first three rows of
rings.
7. The gas turbine engine according to claim 1 wherein the
compressed gas is air.
8. A gas turbine engine comprising a turbine section receiving a
hot working fluid, said turbine section including at least one row
of vanes, at least one row of blades and a plurality of ring
segments forming at least one ring, wherein the ring segments
provide a sealing structure at an end of the blades, said turbine
section further including at least one cooling flow channel
providing a flow of cooling air to one or more of the vanes, the
blades or the ring segments, said cooling flow channel including a
converging-diverging nozzle through which the cooling air flows so
as to create a supersonic flow that reduces the temperature of the
cooling air.
9. The gas turbine engine according to claim 8 wherein the cooling
flow channel includes a wide end section coupled to a converging
section.
10. The gas turbine engine according to claim 9 wherein the cooling
flow channel includes a narrow throat section coupled to the
converging section at one end and a diverging section at an
opposite end.
11. The gas turbine engine according to claim 10 wherein the
cooling flow channel includes a wide output end section at an
output end of the nozzle coupled to the diverging section where the
supersonic flow occurs.
12. The gas turbine engine according to claim 8 wherein the turbine
section includes four rows of vanes, four rows of blades and four
rows of rings.
13. The gas turbine engine according to claim 8 wherein the at
least one cooling flow channel provides a flow of cooling air to
the ring segments.
14. The gas turbine engine according to claim 13 wherein the nozzle
is provided within a plug positioned within the cooling flow
channel.
15. A gas turbine engine comprising: a compressor section being
operable to produce a compressed gas; a combustion section in fluid
communication with the compressor section that receives the
compressed gas, said combustion section mixing the compressed gas
with a fuel and combusting the compressed gas and fuel mixture to
produce a hot working fluid; and a turbine section in fluid
communication with the combustion section, said turbine section
receiving the hot working fluid to produce mechanical power, said
turbine section including at least one row of vanes, at least one
row of blades and a plurality of ring segments forming at least one
ring, wherein the ring segments provide a sealing structure at an
end of the blades, said ring segments and said at least one row of
vanes being mounted to a vane carrier, said vane carrier including
a ring cooling channel for each ring segment that receives a
portion of the compressed gas to cool the ring segments, said
turbine section further including a plug positioned in each ring
cooling channel and having an internal flow channel, said internal
flow channel defining a converging-diverging nozzle including a
wide end section coupled to a converging section, a narrow throat
section coupled to the converging section at one end and a
diverging section at an opposite end and a wide output end section
at an output end of the nozzle coupled to the diverging section,
wherein the compressed air flows at supersonic speeds in the
diverging section that reduces the temperature of the compressed
gas.
16. The gas turbine engine according to claim 15 wherein the
turbine section includes four rows of vanes, four rows of blades
and four rows of rings.
17. The gas turbine engine according to claim 15 wherein the ring
cooling channels are only provided for the first three rows of
rings.
Description
BACKGROUND OF THE INVENTION
[0001] 1. Field of the Invention
[0002] This invention relates generally to a converging-diverging
nozzle for providing reduced temperature cooling air and, more
particularly, to a converging-diverging nozzle positioned within a
cooling fluid flow path in a turbine section of a gas turbine
engine for providing reduced temperature cooling air for cooling
ring segments.
[0003] 2. Discussion of the Related Art
[0004] The worlds energy needs continue to rise which provides a
demand for reliable, affordable, efficient and
environmentally-compatible power generation. Gas turbine engines
are one known machine that provide efficient power, and often have
application for electric generators in a power plant, or engines in
an aircraft or a ship. A typically gas turbine engine includes a
compressor section, a combustion section and a turbine section. The
compressor section provides a compressed air flow to the combustion
section where the air is mixed with a fuel and ignited to generate
a working gas typically having a temperature above 1300.degree. C.
The working gas expands through the turbine section and is guided
across blades by associated vanes. As the working gas passes
through the turbine section, it causes the blades to rotate, which
in turn causes a shaft to rotate, thereby providing mechanical
work.
[0005] The temperature of the working gas is tightly controlled so
that it does not exceed some predetermined temperature for a
particular turbine engine design because to high of a temperature
can damage various parts and components in the turbine section of
the engine. However, it is desirable to allow the temperature of
the working gas to be as high as possible because the higher the
temperature of the working gas, the faster the flow of the gas,
which results in a more efficient operation of the engine.
[0006] In certain gas engine turbine designs, a portion of the
compressed air flow is also used to provide cooling for certain
components in the turbine section, typically the vanes, blades and
ring segments. Thus, the more cooling and/or the more efficient
cooling that can be provided to these components allows the
components to be maintained at a lower temperature, and thus the
higher the temperature of the working gas can be. In other words,
by reducing the temperature of the compressed gas, less compressed
gas is required to maintain the part at the desired temperature,
resulting in a higher working gas temperature and a greater power
and efficiency from the engine. Further, by using less cooling air
at one location in the turbine section, more cooling air can be
used at another location in the turbine section. For example, in
one known turbine engine design, 80% of the compressed air flow is
mixed with the fuel to provide the working gas and 20% of the
compressed air flow is used to cool the turbine section parts. If
less of that cooling air is used at one particular location as a
result of the cooling air being lower in temperature, then more
cooling air can be used at other areas in the turbine section for
increased cooling.
SUMMARY OF THE INVENTION
[0007] In accordance with the teachings of the present invention, a
converging-diverging nozzle is disclosed that has particular
application for providing a cooling air flow to ring segments in a
gas turbine engine. The gas turbine engine includes a compressor
section that produces compressed air and a combustion section in
fluid communication with the compressor section that receives the
compressed air and mixes it with a fuel to generate a hot working
gas. The gas turbine engine also includes a turbine section in
fluid communication with the combustion section that receives the
hot working gas. The turbine section includes at least one row of
vanes, at least one row of blades and a plurality of ring segments
forming at least one ring. The ring segments and the at least one
row of vanes are mounted to a vane carrier, where the vane carrier
includes a cooling flow channel for each of the ring segments that
receives a portion of the compressed air to cool the ring segments.
A plug is provided in each channel and has an internal bore shaped
to define the converging-diverging nozzle through which the
compressed air flows so as to create a supersonic flow that reduces
the temperature of the compressed gas and thus provides more
cooling for the same amount of air flow.
[0008] Additional features of the present invention will become
apparent from the following description and appended claims, taken
in conjunction with the accompanying drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] FIG. 1 is a cut-away, isometric view of a gas turbine
engine;
[0010] FIG. 2 is a cross-sectional, broken-away view of a portion
of the turbine section of the gas turbine engine shown in FIG.
1;
[0011] FIG. 3 is an enlarged view of a portion of the turbine
section shown in FIG. 2 specifically showing a ring segment;
[0012] FIG. 4 is an isometric view of a cooling channel plug
including a cooling gas flow channel therein;
[0013] FIG. 5 is a cross-sectional view of a converging-diverging
nozzle; and
[0014] FIG. 6 is a graph with entropy on the horizontal axis and
temperature on the vertical axis showing a temperature and pressure
relationship of a gas flowing through the converging-diverging
nozzle.
DETAILED DESCRIPTION OF THE EMBODIMENTS
[0015] The following discussion of the embodiments of the invention
directed to a converging-diverging nozzle for providing a reduced
temperature cooling gas flow is merely exemplary in nature and is
in no way intended to limit the invention or its applications or
uses. For example, the discussion below states that the
converging-diverging nozzle has particular application for
providing a cooling gas flow to ring segments in the turbine
section of a gas turbine engine. However, as will be appreciated by
those skilled in the art, the converging-diverging nozzle of the
invention may have application for providing a cooling gas flow to
other components in a gas turbine engine, and other cooling gas
flow applications other than for gas turbine engines.
[0016] FIG. 1 is a cut-away isometric view of a gas turbine engine
10 including a compressor section 12, a combustion section 14 and a
turbine section 16, where operation of the engine 10 causes a
central shaft 18 or rotor to rotate, thus creating mechanical
power. Rotation of the shaft 18 draws air into the compressor
section 12 where it is directed by vanes 22 and compressed by
rotating blades 20 to be delivered to the combustion section 14
where the compressed air is mixed with ignition fuel, such as
propane, and where the fuel/air mixture is ignited to create a hot
working gas. More specifically, the combustion section 14 includes
a number of circumferentially disposed combustion chambers 26 each
receiving the ignition fuel that is sprayed into the chamber 26 by
an injector (not shown) and mixed with the compressed air to be
combusted to create the working gas, which is directed by a nozzle
28 into the turbine section 16. The working gas is directed by
circumferentially disposed stationary vanes 32 (see FIG. 2) in the
turbine section 16 to flow across circumferentially disposed
rotatable turbine blades 34, which causes the turbine blades 34 to
rotate, thus rotating the shaft 18. Each group of the
circumferentially disposed stationary vanes 32 defines a row of the
vanes 32 and each group of the circumferentially disposed blades 34
defines a row of the blades 34. Once the working gas passes through
the turbine section 16 it is output from the engine 10 as an
exhaust gas through output nozzle 36.
[0017] In this non-limiting embodiment, the turbine section 16
includes four rows 38 of the rotating blades 34 and four rows of
the stationary vanes 32 in an alternating sequence. In other gas
turbine engine designs, the turbine section 16 may include more or
less rows of the turbine blades 34. It is noted that the most
forward row of the turbine blades 34 and the vanes 32 receive the
highest temperature of the working gas, where the temperature of
the working gas decreases as it flows through the turbine section
16.
[0018] FIG. 2 is a cross-sectional type view of a portion of the
turbine section 16 and specifically a turbine vane carrier 40. The
vane carrier 40 includes a number of mounting structures 42 for
securing each of the plurality of vanes 32 for each of the four
rows of the vanes 32 to the vane carrier 40. A plurality of
circumferentially disposed ring segments 46 are mounted to the vane
carrier 40 by a separate mounting structure 44 and define a ring
where the ring segments 46 for a particular ring are positioned
adjacent to each other to form the ring, and where a separate ring
is provided for each row of the blades 34. As is well understood by
those skilled in the art, the ring segments 46 provide a sealing
structure that allows the blades 34 (not shown in FIG. 2) to rotate
in close proximity to the ring segments 46 to limit the amount of
the working gas that can flow past the blades 34 without causing
the blades 34 to rotate. The number and size of the ring segments
46 will be different for each blade row or stage, and would be
different from turbine design to turbine design.
[0019] FIG. 3 is a blown-up view of one of the ring segments 46
provided between the first and second rows of the vanes 32 to
illustrate how the ring segments 46 are cooled using the compressed
air from the compression section 12 during operation of the engine
10. Particularly, a cylindrical channel 50 is formed through the
vane carrier 40 for each ring segment 46 in each ring from an
outside surface of the vane carrier 40 to the location where the
ring segment 46 is mounted to the vane carrier 40. In this
particular non-limiting design, the channel 50 includes a wide top
portion 52, a narrow middle portion 54 and a wide bottom portion
56. A plug 60 is mounted within the channel 50 so that a head 62 of
the plug 60 sits in the top portion 52 and a body 64 of the plug 60
extends into the narrow portion 54. In this particular design, only
the ring segments 46 in the first three rows of the rings receive
the cooling air as it is not necessary to cool the last row of the
ring segments 46 because the temperature of the working gas is
reduced to a low enough temperature by that location in the turbine
section 16.
[0020] FIG. 4 is an isometric view of the plug 60 separated from
the vane carrier 40. A cylindrical metering channel 66 is formed
axially through the plug 60 so that a controlled amount of the
cooling air can flow through the plug 60 and into the lower portion
of the channel 50. Suitable plumbing (not shown) is provided, such
as appropriately formed channels in the vane carrier 40 and other
structural elements of the turbine section 16, to allow the
compressed cooling air flow from the compressor section 12 to be
directed through the plug 60 and into the channel 50 to cool the
ring segment 64. The size of the channel 50 in the vane carrier 40
and the size of the channel 64 in the plug 60 are carefully
engineered so that the desired amount of mass air flow is provided
to all of the ring segments 46 in all of the rings, and be able to
provide a cooling air flow to the other parts in the turbine
section 16 without compromising turbine engine efficiency.
[0021] The present invention proposes replacing the plug 60 with a
similarly configured plug of the same size that would fit in the
existing cooling channel 50 for the ring segment 46 for a
particular turbine engine and which includes an internal flow
channel configured as a converging-diverging nozzle. FIG. 5 is a
cross-sectional view of a converging-diverging nozzle 70 having a
flow channel 72 suitable for this purpose. Particularly, the nozzle
70 would replace the cylindrical channel 66 in the plug 60 through
which the compressed air flows. As will be discussed in more detail
below, the channel 72 has a particular hour-glass shape that
provides subsonic-sonic-supersonic flow that causes a decrease in
temperature of the air as it flows through the channel 72. The
physics that causes changes in temperature and pressure of gas
flowing through a converging-diverging nozzle is well understood by
those skilled in the art.
[0022] The flow channel 72 of the nozzle 70 has the shape shown in
FIG. 5, where the gas flows through the channel 72 from an input
end 74 to an output end 76. The channel 72 includes a wide input
section 78 that is coupled to a converging section 80 which in turn
is coupled to a narrow throat section 82 representing the narrowest
part of the channel 72. The throat section 82 is coupled to a
diverging section 84 that is coupled to an annular shoulder 86 that
defines a wide end section 88 at the output end 76 of the nozzle 70
having the same diameter as the wide end section 88.
[0023] The mass flow rate of the compressed air provided to the
plug 60 is such that sufficient cooling is provided during
operation of the engine 10. Further, as the air flows through the
channel 72 the entropy s of the air also is maintained relatively
constant as the pressure and temperature of the air flow changes.
The velocity of the air flow increases as it flows through the
converging flow channel 72 and correspondingly, the pressure of the
air flow and temperature of the air flow decreases. This
relationship between pressure and temperature of the air flow
through the channel 72 is illustrated in the graph of FIG. 6 with
entropy s on the horizontal axis and temperature T on the vertical
axis. At the input end 74 of the nozzle 70, the pressure of the gas
is represented by P.sub.o, which has a certain temperature T.sub.o.
As the air flow converges through the converging section 80 its
pressure decreases as the velocity of the gas increases and its
temperature decreases, where the pressure is represented by P.sub.i
and the temperature is represented by T.sub.i at this location in
the channel 72. The design of the nozzle 70 is such that for a
particular mass air flow rate, the air flow is subsonic through the
converging section 80 and reaches sonic speeds at the throat
section 82. When the gas flows into the diverging section 84, the
air flow accelerates to supersonic flow, further reducing the
pressure and temperature of the gas represented by pressure P.sub.e
and T.sub.e, respectively. When the gas enters the section 88 of
the nozzle 70, it is flowing supersonically, which causes it to
have a reduced pressure P.sub.b and a reduced temperature
T.sub.b.
[0024] By providing the same amount of compressed air to the plug
60 at the same flow rate, the nozzle 70 will provide that air flow
to the ring segment 46 at a lower temperature than the channel 66,
which provides increased cooling to the ring segment 46. By
providing the lower temperature air to the ring segment 46, a
number of design changes can be made to the turbine section 16 to
increase the efficiency of the engine 10. For example, if the same
amount of the compressed air is provided to the turbine section 16
for cooling, then a reduced percentage of that air can be provided
to cool the ring segments 16 and achieve the same amount of cooling
as a result of the decrease in temperature of the cooling air. The
extra cooling air now available because of the reduced amount of
compressed air provided to the ring segments 46 can be used to
provide increased cooling to other components in the turbine
section 16, such as the vanes 32 and the blades 34. By providing
increased cooling to these parts of the turbine section 16, the
turbine inlet temperature can be increased, which increases the
power and efficiency of the turbine engine 10. Alternately, by
providing a reduced amount of cooling air to the ring segments 46
because of the decrease in temperature provided by the nozzle 70,
the turbine design can be changed where a lower percentage of the
compressed air is provided for cooling and more air is provided for
combustion, which can be used to lower NOx emissions.
[0025] It is noted that the shape of the nozzle 70 is a general
shape for a particular application for cooling the ring segments
46. The flow rate and pressure provided to each of the ring
segments 46 in each of the rows is different, which would require a
slightly different shape to the nozzle to achieve the desired
supersonic flow. Thus, all of the diameters and configurations of
the sections in the nozzle are tightly engineered to get the flow
and pressure effect to create the supersonic shock wave to reduce
the temperature of the air.
[0026] The foregoing discussion discloses and describes merely
exemplary embodiments of the present invention. One skilled in the
art will readily recognize from such discussion, and from the
accompanying drawings and claims, that various changes,
modifications and variations can be made therein without departing
from the scope of the invention as defined in the following
claims.
* * * * *